Referring to FIG. 1, a turbine engine 10 can generally include a compressor section 12, a combustor section 14 and a turbine section 16. A centrally disposed rotor 18 can extend through the three sections. Portions of the rotor 18 can be protected by a rotor shaft cover 19.
The turbine section 16 can include alternating rows of vanes 20 and rotating blades 22. Each row of blades 22 can include a plurality of airfoils attached to a disc 24 provided on the rotor 18. The rotor 18 can include a plurality of axially-spaced discs 24. The blades 22 can extend radially outward from the discs 24.
Each row of vanes 20 can be formed by attaching a plurality of vanes 20 to the stationary support structure in the turbine section 16. For instance, the vanes 20 can be mounted on a vane carrier 26 that is attached to the outer casing 28. The vanes 20 can extend radially inward from the vane carrier 26.
In operation, the compressor section 12 can induct ambient air and can compress it. The compressed air 32 from the compressor section 12 can enter a chamber 34 enclosing the combustor section 14. The compressed air 32 can then be distributed to each of the combustors 36 (only one of which is shown). In each combustor 36, the compressed air 32 can be mixed with the fuel. The air-fuel mixture can be burned to form a hot working gas 38. The hot gas 38 can be routed to the turbine section 16 by a duct 42, sometimes referred to as a transition. As it travels through the rows of vanes 20 and blades 22, the gas 38 can expand and generate power that can drive the rotor 18. The expanded gas 40 can then be exhausted from the turbine 16.
During engine operation, the axial and radial displacement of an inner support structure 44 (which can be affected by at least the displacement of the vane carrier 26 and the outer casing 28) and a support structure 46 (which can be affected by displacement of at least the shaft cover 19) at the outlet of the transition duct 42 is not the same due to differential thermal growth and movement of these structures. As a result, there can be relatively large relative movements between these structures 44, 46.
These relative movements can produce high stresses within the first row of vanes. In addition, these relative movements can cause ID-to-OD rocking of the vane between the inner platform element of the vane and the transition duct from the combustor, potentially resulting in substantial gas leakage and loss of efficiency due to the large relative displacement.
One system for addressing such issues in connection with conventional transition duct systems is described in U.S. Patent Application Publication No. 20080008584. In such systems, exhaust flow is directed straight in the axially aft direction. There is a minimum amount of space available between transition ducts to receive a support member. However, there is not enough space for a substantial support member because any support member must be circumferentially narrow to fit in the limited available space between neighboring transitions.
Moreover, such a system cannot be applied to transition ducts that are configured to eliminate the first stage row of turbine vanes. An example of such a transition duct system is described in U.S. Patent Application Publication No. 20070017225. In such systems, the transitions generally supply combustion gases with high tangential velocity directly to a first row of blades. However, such a configuration results in the outlets of the transition ducts being arranged so close together that there is no longer room to physically fit any support members, thereby leaving such systems prone to the transition duct outlet displacement issues similar to those that have plagued systems in the past. Thus, there is a need for a system that can minimize these concerns.